JP2017210959A - Cooling passage for gas turbine rotor blade - Google Patents

Abstract

PROBLEM TO BE SOLVED: To provide a cooling passage for a gas turbine rotor blade.SOLUTION: The present disclosure is directed to a rotor blade 100 for a gas turbine engine. The rotor blade 100 includes an airfoil 126, a tip shroud 128 having side surfaces 146, 148 and a radially outer surface 142, and a transition portion 150 coupling the tip shroud 128 to the airfoil 126. The airfoil 126, the transition portion 150, and the tip shroud 128 collectively define primary cooling passages 152A, 152B therein. The primary cooling passages 152A, 152B include primary cooling passage outlets 154A, 154B defined by the side surfaces 146, 148 of the tip shroud 128.SELECTED DRAWING: Figure 5

Description

  The present disclosure relates generally to gas turbine engines. More particularly, the present disclosure relates to gas turbine engine rotor blades.

  A gas turbine engine typically includes a compressor section, a combustion section, a turbine section, and an exhaust section. The compressor section gradually increases the pressure of the working fluid flowing into the gas turbine engine, and supplies the compressed working fluid to the combustion section. Compressed working fluid and fuel (eg, natural gas) are mixed in the combustion section and burned in the combustion chamber to produce high pressure and high temperature combustion gases. Combustion gas flows from the combustion section to the turbine section where it expands to produce work. For example, the expansion of the combustion gas in the turbine section can generate electricity by rotating a rotor shaft connected to a generator, for example. Next, the combustion gas is exhausted from the gas turbine through the exhaust section.

  The turbine section includes a plurality of rotor blades that extract kinetic and / or thermal energy from the combustion gases flowing therethrough. In certain embodiments, some or all of the plurality of rotor blades include a tip shroud coupled to the airfoil portion by a fillet portion. These rotor blades generally operate in extremely hot environments. As such, the rotor blade typically includes one or more cooling passages defined therein. During operation of a gas turbine engine, a cooling medium, such as compressed air, flows through one or more cooling passages to cool the rotor blades. Nevertheless, conventional cooling passage configurations that provide sufficient cooling to the fillet portion and tip shroud increase the weight of the rotor blade, which may be undesirable.

U.S. Pat. No. 8,348,612

  Aspects and advantages of the present technology are set forth in part in the following description, or are obvious from the description, or can be learned by practice of the technology.

  In one aspect, the present disclosure is directed to a gas turbine engine rotor blade. The rotor blade includes an airfoil, a tip shroud having a side surface and a radially outer surface, and a transition portion coupling the tip shroud to the airfoil. The airfoil, transition portion, and tip shroud collectively define a primary cooling passage therein. The primary cooling passage includes a primary cooling passage outlet defined by a side surface of the tip shroud.

  A further aspect of the present disclosure relates to a gas turbine engine having a compressor portion, a combustion portion, and a turbine portion. The turbine portion includes one or more rotor blades. Each rotor blade includes an airfoil, a tip shroud having sides and a radially outer surface, and a transition portion coupling the tip shroud to the airfoil. The airfoil, transition portion, and tip shroud collectively define a primary cooling passage therein. The primary cooling passage includes a primary cooling passage outlet defined by a side surface of the tip shroud.

  These and other features, aspects and advantages of the present technology will be better understood with reference to the following description and appended claims. The accompanying drawings, which are incorporated in and constitute a part of this specification, illustrate embodiments of the technology and, together with the description, serve to explain the principles of the technology.

  The complete and possible disclosure of the present technology, including its best mode, is directed to those skilled in the art and is described herein, which refers to the following accompanying drawings.

1 is a schematic diagram of an exemplary gas turbine engine that may incorporate various embodiments disclosed herein. FIG. 2 is a front view of an exemplary rotor blade that may be incorporated into the gas turbine shown in FIG. 1 according to embodiments disclosed herein. FIG. FIG. 3 is a top view of the exemplary rotor blade shown in FIG. 2, further illustrating various features thereof. FIG. 4 is an enlarged perspective view of a portion of the rotor blade shown in FIGS. 2 and 3, showing the radially outer portion of the airfoil and tip shroud. FIG. 5 is a cross-sectional view of the airfoil and tip shroud taken generally along line 5-5 of FIG. 4, illustrating one embodiment of a first primary cooling passage and a second primary cooling passage. FIG. 5 is a cross-sectional view of the airfoil and tip shroud taken generally along line 5-5 of FIG. 4, showing another embodiment of the first primary cooling passage. FIG. 4 is a front view of a portion of the tip shroud, showing a first primary cooling passage outlet. 2 is an enlarged cross-sectional view of a first primary cooling passage showing one embodiment of one or more turbulators. FIG. FIG. 6 is an enlarged cross-sectional view of a first primary cooling passage showing an alternative embodiment of one or more turbulators. FIG. 6 is an enlarged cross-sectional view of a first primary cooling passage showing a further embodiment of one or more turbulators. FIG. 6 is a cross-sectional view of one embodiment of a jacketed core and mold used to form first and / or second primary cooling passages. FIG. 5 is a cross-sectional view of the airfoil and tip shroud taken generally along line 5-5 of FIG. 4, showing a further embodiment of the first primary cooling passage.

  Repeat use of reference characters in the present specification and drawings is intended to represent same or analogous features or elements of the present technology.

  Reference will now be made in detail to the embodiments of the technology, one or more examples of which are illustrated in the accompanying drawings. The detailed description uses numerical and letter designations to refer to features of the drawings. Like or similar symbols in the drawings and description are used to refer to like or similar parts of the technology. In this specification, the terms "first", "second", and "third" can be used interchangeably to distinguish one component from another component; It is not intended to indicate the location or importance of individual components. The terms “upstream” and “downstream” indicate relative directions for fluid flow in the fluid path. For example, “upstream” indicates the direction in which the fluid flows therefrom, and “downstream” indicates the direction in which the fluid flows.

  Each example is provided by way of explanation of the present technology and not limitation of the present technology. In fact, it will be apparent to those skilled in the art that modifications and variations can be made in the present invention without departing from the scope or spirit of the technology. For example, features illustrated or described as part of one embodiment can be used in another embodiment to provide a still further embodiment. Accordingly, the technology is intended to embrace such modifications and variations that fall within the scope of the appended claims and their equivalents. Although industrial or terrestrial gas turbines are shown and described herein, the techniques shown and described herein are intended for terrestrial and / or industrial use unless otherwise specified in the claims. It is not limited to gas turbines. For example, the techniques described herein may be used with any type of turbine, including but not limited to aviation gas turbines (eg, turbofans, etc.), steam turbines, and marine gas turbines.

  Referring now to the drawings, wherein like numerals indicate like elements throughout the drawings, FIG. 1 schematically illustrates a gas turbine engine 10. It should be understood that the turbine engine 10 of the present disclosure need not be a gas turbine engine and may be any suitable turbine engine, such as a steam turbine engine or other suitable engine. The gas turbine engine 10 includes an inlet portion 12, a compressor portion 14, a combustion portion 16, a turbine portion 18, and an exhaust portion 20. The compressor portion 14 and the turbine portion 18 may be coupled by a shaft 22. The shaft 22 may be a single shaft or a plurality of shaft segments coupled together to form the shaft 22.

  The turbine section 18 generally includes a rotor shaft 24 having a plurality of rotor disks 26 (one of which is shown) and a plurality of interconnected to the rotor disks 26 extending radially outward from the rotor disks 26. Rotor blade 28. Each rotor disk 26 may be coupled to a portion of the rotor shaft 24 that extends through the turbine section 18. The turbine section 18 further includes an outer casing 30 that circumferentially surrounds the rotor shaft 24 and the rotor blades 28, thereby at least partially defining a hot gas path 32 through the turbine section 18.

  During operation, air or another working fluid flows through the inlet section 12 to the compressor section 14 where the air is gradually compressed and compressed air is supplied to a combustor (not shown) in the combustion section 16. Provided. The pressurized air is mixed with fuel and burned in each combustor to produce combustion gas 34. The combustion gas 34 flows along the hot gas path 32 from the combustion section 16 to the turbine section 18 where energy (motion and / or heat) is transferred from the combustion gas 34 to the rotor blade 28 and rotates the rotor shaft 24. The mechanical rotational energy can then be used to power the compressor section 14 and / or generate electricity. The combustion gas 34 exiting from the turbine unit 18 is exhausted from the gas turbine engine 10 via the exhaust unit 20.

  2 and 3 can incorporate one or more embodiments disclosed herein and can be incorporated into the turbine section 18 of the gas turbine engine 10 in place of the rotor blade 28 shown in FIG. 1 is a diagram of an exemplary rotor blade 100. FIG. As shown in FIGS. 2 and 3, the rotor blade 100 defines an axial direction A, a radial direction R, and a circumferential direction C. The radial direction R extends substantially perpendicular to the axial direction A, and the circumferential direction C extends substantially concentrically around the axial direction A.

  As shown in FIGS. 2 and 3, the rotor blade 100 includes a platform 102 that generally serves as a radially inner flow boundary for the combustion gas 34 flowing through the hot gas path 32 of the turbine section 18 (FIG. 1). More specifically, platform 102 includes a radially inner surface 104 that is radially spaced from radially outer surface 106. The platform 102 also includes a leading edge 108 that is axially spaced from the trailing edge 110. The leading edge 108 is disposed in the flow of combustion gas 34 and the trailing edge 110 is disposed downstream of the leading edge 108. Further, the platform 102 includes a pressure side slash surface 112 circumferentially spaced from the suction side slash surface 114.

  As shown in FIG. 2, the rotor blade 100 includes a shank portion 116 that extends radially inward from the radially inner surface 104 of the platform 102. One or more angel wings 118 may extend axially outward from the shank portion 116. The shank portion 116 and platform 102 collectively define the shank pocket 120 in the embodiment shown in FIG. Nevertheless, in some embodiments, the shank portion 116 and the platform 102 may not define the shank pocket 120.

  The rotor blade 100 also includes a root portion 122 that extends radially inward from the shank portion 116. The root portion 122 can interconnect or secure the rotor blade 100 to the rotor disk 26 (FIG. 1). In the embodiment shown in FIG. 2, the root portion 122 has a fir tree configuration. Nevertheless, the root portion 122 may have any suitable configuration (eg, a dovetail configuration). Further, the root portion 122 can define an intake port 124 that allows cooling air to enter the rotor blade 100.

  The rotor blade 100 further includes an airfoil 126 that extends radially outward from the radially outer surface 106 of the platform 102 to the tip shroud 128. As such, the tip shroud 128 can generally define the radially outermost portion of the rotor blade 100. The airfoil 126 is coupled to the platform 102 at the airfoil root 130 (ie, the intersection between the airfoil 126 and the platform 102). In some embodiments, the airfoil root 130 may be a radius or fillet (not shown) that transitions between the airfoil 126 and the platform 102. In this regard, the airfoil 126 defines an airfoil span 132 that extends between the airfoil root 130 and the tip shroud 128. The airfoil 126 also includes a pressure side wall 134 and an opposing suction side wall 136. The pressure side wall 134 and the pressure side wall 136 are joined or interconnected together at the leading edge 138 of the airfoil 126 oriented in the flow of the combustion gas 34. The pressure side wall 134 and the pressure side wall 136 are also joined or interconnected together at the trailing edge 140 of the airfoil 126 spaced downstream of the leading edge 138. The pressure side wall 134 and the suction side wall 136 are continuous around the leading edge 138 and the trailing edge 140. The pressure side wall 134 is generally concave and the pressure side wall 136 is generally convex.

  As described above, the tip shroud 128 is disposed at the radially outer end of the rotor blade 100. The tip shroud 128 reduces the amount of combustion gas 34 that escapes beyond the rotor blade 100. In the embodiment shown in FIGS. 3 and 4, the tip shroud 128 includes a radially outer surface 142 having rails 144 extending radially outward therefrom. Alternative embodiments may include more rails 144 (eg, two rails 144, three rails 144, etc.) or no rails 144 at all. The tip shroud 128 also includes a pressure side 146 and a suction side 148 interconnected by the leading edge 138 and trailing edge 140 of the airfoil 126. In the embodiment shown in FIG. 3, the pressure side surface 146 and the suction side surface 148 are generally not aligned with the pressure side wall 134 and the suction side wall 136 of the airfoil 126. In addition, the tip shroud 128 includes a radially inner surface 178 disposed radially inward from its radially outer surface 142.

  Referring now to FIG. 5, the transition portion 150 couples the tip shroud 128 to the radially outer end of the airfoil 126. The transition portion 150 includes an outer surface 180 that is exposed to the combustion gas 34 in the hot gas path 32. The outer surface 180 couples the outer surfaces of the pressure and suction walls 134, 136 of the airfoil 126 to the radially inner surface 178 of the tip shroud 128. In the embodiment shown in FIGS. 5 and 6, transition portion 150 is a fillet portion that transitions between airfoil 126 and tip shroud 128. In another embodiment, transition portion 150 may include a chamfer (not shown) or other suitable transition between airfoil 126 and tip shroud 128.

  The rotor blade 100 defines therein one or more cooling primary passages through which cooling air flows. In the embodiment shown in FIG. 5, airfoil 126, transition portion 150, and tip shroud 128 collectively define a first primary cooling passage 152A and a second primary cooling passage 152B therein. Nevertheless, the rotor blade 100 can define more or fewer primary cooling passages as needed or desired. In fact, the rotor blade 100 can define any number of primary cooling passages as long as it defines at least one primary cooling passage.

  Each of the first and second primary cooling passages 152A, 152B includes a corresponding inlet. In some embodiments, the intake port 124 (FIG. 2) defined by the root portion 122 may be the inlet of the first and second primary cooling passages 152A, 152B. In such an embodiment, root portion 122, shank portion 116, platform 102, airfoil 126, transition portion 150, and tip shroud 128 collect therein first and second primary cooling passages 152A, 152B. To be defined. In other embodiments, the inlets of the first and second primary cooling passages 152A, 152B are chambers, passages, or cavities defined by the root portion 122, the shank portion 116, the platform 102, and / or the airfoil 126. (E.g., shank pocket 120).

  The first and second primary cooling passages 152A and 152B terminate at a first primary cooling passage outlet 154A and a second primary cooling passage outlet 154B, respectively. Both the first and second primary cooling passage outlets 154A, 154B are defined by a tip shroud 128. In the embodiment shown in FIG. 5, the pressure side 146 of the tip shroud 128 defines a first primary cooling passage outlet 154A and the suction side 148 of the tip shroud 128 defines a second primary cooling passage outlet 154B. Nevertheless, each of the first and second primary cooling passage outlets 154A, 154B may be defined by either the pressure side 146 or the suction side 148 of the tip shroud 128.

  As described above, airfoil 126, transition portion 150, and tip shroud 128 collectively define first and second primary cooling passages 152A, 152B therein. More specifically, the airfoil 126 defines an airfoil portion 156A of the first primary cooling passage 152A and an airfoil portion 156B of the second primary cooling passage 152B. The transition portion 150 of the rotor blade 100 defines a transition portion 158A of the first primary cooling passage 152A and a transition portion 158B of the second primary cooling passage 152B. Preferably, the transition portions 158A, 158B should be located proximate to the outer surface 180 of the transition portion 150 of the rotor blade 100 so as to improve cooling of the rotor blade 100. That is, the closer the transition portions 158A, 158B of the first and second primary cooling passages 152A, 152B are to the outer surface 180 of the transition portion 150, the better the cooling. The tip shroud 128 defines a tip shroud portion 160A of the first primary cooling passage 152A and a tip shroud portion 160B of the second primary cooling passage 152B.

  5 and 6 show different configurations of the airfoil portions 156A, 156B, transition portions 158A, 158B and tip shroud portions 160A, 160B of the first and second primary cooling passages 152A, 152B. FIG. 6 omits the second primary cooling passage 152B for clarity. In the embodiment shown in FIG. 5, the airfoil portions 156A, 156B of the first and second primary cooling passages 152A, 152B are straight and the transition portions 158A of the first and second primary cooling passages 152A, 152B, 158B and tip shroud portions 160A, 160B are curved. However, in the embodiment shown in FIG. 6, the airfoil portion 156A and transition portion 158A of the first primary cooling passage 152A are curved and the tip shroud portion 160A is straight. Nevertheless, any of the airfoil portions 156A, 156B, transition portions 158A, 158B, and tip shroud portions 160A, 160B of the first and second primary cooling passages 152A, 152B are straight, curved, or Any combination thereof may be used.

  As described above, in the embodiment shown in FIG. 5, the transition portion 150 of the rotor blade 100 has a fillet portion, and the second transition portion 158B of the second primary cooling passage 152B is curved. In this embodiment, the curvature of the second transition portion 158B of the second primary cooling passage 152B substantially follows the curvature of the fillet portion. That is, the transition portion 158B of the second primary cooling passage 152B is uniformly spaced from the outer surface of the fillet portion along the entire length of the fillet portion. The curvature of the first transition portion 158A of the first primary cooling passage 152A may also substantially follow the curvature of the fillet portion. Nevertheless, in some embodiments, the curvature of the first and second transitional portions 158A, 158B of the first and second primary cooling passages 152A, 152B may not follow the curvature of the fillet portion. Good.

  The first and second primary cooling passages 152A, 152B and the corresponding first and second primary cooling passage outlets 154A, 154B can have any suitable cross-sectional shape. In one embodiment, the first and second primary cooling passages 152A, 152B and the corresponding first and second primary cooling passage outlets 154A, 154B have a circular cross-sectional shape. In other embodiments, the first and second primary cooling passages 152A, 152B and the corresponding first and second primary cooling passage outlets 154A, 154B have an oval or elliptical shape as shown in FIG. You may have. If the corresponding first and / or second primary cooling passage outlets 154A, 154B have an oval cross section, the pressure side and / or suction side surfaces 146, 148 are thinner (ie shorter in the radial direction R). ) Nevertheless, the first and second primary cooling passages 152A, 152B and the corresponding first and second primary cooling passage outlets 154A, 154B may have any suitable cross-sectional shape. The first primary cooling passage 152A and the first primary cooling passage outlet 154A may have the same or different cross-sectional shapes as the second primary cooling passage 152B and the second primary cooling passage outlet 154B. In some embodiments, the first and second primary cooling passages 152A, 152B may have a different cross-sectional shape than the corresponding first and second primary cooling passage outlets 154A, 154B. The first and second primary cooling passages 152A, 152B preferably have a constant diameter. Nevertheless, the diameter of the first and / or second primary cooling passages 152A, 152B may vary along its length. In the embodiment shown in FIG. 5, for example, the first primary cooling passage 152 </ b> A has a first diameter 168 that is smaller than the second diameter 170.

  Each of the first and second primary cooling passages 152A, 152B may include one or more turbulators disposed therein. The one or more turbulators create turbulence in the cooling air flowing through the corresponding primary cooling passages 152A, 152B, thereby increasing the heat transfer rate between the rotor blade 100 and the cooling air. 8A-8C show various embodiments of turbulators. For example, the one or more turbulators may include one or more rectangular protrusions 172 (FIG. 8A), one or more hemispherical protrusions 174 (FIG. 8B), one or more dimples 176 (FIG. 8C), or these Any combination of the above may be used. Indeed, the one or more turbulators may be any suitable feature disposed in the first and / or second primary cooling passages 152A, 152B that create turbulence therein.

  In some embodiments, each of the first and second primary cooling passages 152A, 152B can be formed using a jacketed core 182. As shown in FIG. 9, the jacketed core 182 includes an annular sleeve 184 filled with a core material 186. In some embodiments, the annular sleeve 184 is formed from nickel or a nickel-based alloy and the core material 186 is a refractory ceramic (eg, silica, alumina, mullite, etc.). The jacketed core 182 can be modified into a configuration having various straight and / or curved portions, such as the shape of the first and second primary cooling passages 152A, 152B. The cross-sectional shape of the jacketed core 182 may also be modified to form the oval cross section shown in FIG. Once formed into the desired shape of the first or second primary cooling passages 152A, 152B, the jacketed core 182 is disposed within a mold 188 for forming the rotor blade 100. The mold 188 is then filled with a molten material 190 (eg, a molten nickel-based superalloy) that is used to form the rotor blade 100. During the casting process, molten material 190 poured into mold 188 absorbs annular sleeve 184. In this regard, the materials forming the annular sleeve 184 and the rotor blade 100 are thoroughly mixed so that there is no separate boundary between them upon completion of the casting process. FIG. 9 shows the jacketed core 182 after the molten material 190 has been injected into the mold 188 and before the molten material 190 absorbs the annular sleeve 184. The core material 186 is then removed from the rotor blade 100 after casting (eg, by chemical leaching), leaving a primary cooling passage. However, in alternative embodiments, the first and second primary cooling passages 152A, 152B may be formed using any suitable method or process.

  As described above, the first and second primary cooling passages 152A, 152B guide cooling air through the rotor blade 100 and cool various portions thereof. More specifically, the cooling air (for example, air discharged from the compressor unit 14 (FIG. 1)) passes through the intake port 124 (FIG. 2), for example, and the first and second primary cooling passages 152A and 152B. to go into. Cooling air flows through at least the airfoil portions 156A, 156B, transition portions 158A, 158B, and tip shroud portions 160A, 160B of the first and second primary cooling passages 152A, 152B. Depending on the location of the inlet, the cooling air may flow through other parts of the rotor blade 100. For example, if the intake port 124 is the inlet of the first and / or second primary cooling passages 152A, 152B, the cooling air will flow to the root portion 122 (FIG. 2), the shank portion 116 (FIG. 2), and the platform It may flow through 102 (FIG. 2). While flowing through the first and second primary cooling passages 152A, 152B, the cooling air absorbs heat from the rotor blade 100 and cools it. That is, the heat from the rotor blade 100 is transferred to the cooling air by convection. Cooling air exits the first and second primary cooling passages 152A, 152B through corresponding first and second primary cooling passage outlets 154A, 154B and flows into the hot gas path 32 (FIG. 1).

  FIG. 10 illustrates an alternative embodiment of the rotor blade 100. In this embodiment, rotor blade 100 defines therein one or more secondary cooling passages in fluid communication with first primary cooling passage 152A. As described in more detail below, the one or more secondary cooling passages may be in addition to or in place of the convective cooling provided by the first and second primary cooling passages 152A, 152B. 100 provides membrane cooling and / or additional convective cooling.

  In the embodiment shown in FIG. 10, the tip shroud 128 of the rotor blade 100 defines a first secondary cooling passage 162A and a second secondary cooling passage 162B therein. The tip shroud 128 and the transition portion 150 of the rotor blade 100 collectively define a third secondary cooling passage 162C and a fourth secondary cooling passage 162D therein. In other embodiments, the rotor blade 100 can define more or fewer secondary cooling passages as needed or desired. Although FIG. 10 shows only the first primary cooling passage 152A, additional secondary cooling passages may include other primary cooling passages defined by the rotor blade 100 (eg, the second primary cooling passage 152B). It may be in fluid communication.

  As shown in FIG. 10, the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, and 162D are respectively connected to the first secondary cooling passage from the first primary cooling passage 152A. Extends to outlet 164A, second secondary cooling passage outlet 164B, third secondary cooling passage outlet 164C, and fourth secondary cooling passage outlet 164D. In the embodiment shown in FIG. 10, the radially outer surface 142 of the tip shroud 128 defines first and second secondary cooling passage outlets 164A, 164B, and the transition portion 150 includes third and fourth secondary cooling. A passage outlet 164C, 164D is defined. In other embodiments, one of the transition portion 150 or the radially outer surface 142 of the tip shroud 128 defines all of the first, second, third, and fourth outlets 164A, 164B, 164C, 164D. can do. Nevertheless, the first, second, third, and fourth outlets 164A, 164B, 164C, 164D may be defined by any surface of the transition portion 150 or the tip shroud 128.

  The first, second, third, and fourth secondary cooling passages 162 </ b> A, 162 </ b> B, 162 </ b> C, 162 </ b> D can be oriented to direct cooling air to the outer surface of the rotor blade 100. More specifically, the first and second secondary cooling passages 162A, 162B can direct air to the radially outer surface 142 of the tip shroud 128. In this regard, the first and second secondary cooling passages 162A, 162B are preferably oriented at an angle with respect to the radially outer surface 142 of the tip shroud 128. The third and fourth secondary cooling passages 162 </ b> C, 162 </ b> D can guide air to the outer surface 180 of the transition portion 150. Accordingly, the third and fourth secondary cooling passages 162C, 162D may be oriented at an angle with respect to the outer surface 180 of the transition portion 150. Indeed, the third and fourth secondary cooling passages 162C, 162D may be oriented tangential to the outer surface 180 of the transition portion 150. Preferably, the cooling air exits the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D that are substantially parallel to the outer surface of the rotor blade 100. By orienting the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D and directing cooling air to the outer surface of the rotor blade 100, film cooling is facilitated. In alternative embodiments, the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D may not direct cooling air to the outer surface of the rotor blade 100.

  The plug 166 may be disposed at the first primary cooling passage outlet 154A of the first primary cooling passage 152A in embodiments including one or more secondary cooling passages. More specifically, the plug 166 blocks the flow of cooling air passing through the first primary cooling passage outlet 154A. In this way, all of the cooling air present in the first primary cooling passage 152A flows into the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D. The cooling air then exits the rotor blade 100 through the corresponding first, second, third, or fourth secondary cooling passage outlets 164A, 164B, 164C, 164D. In this regard, the first primary cooling passage 152A acts as a plenum when occluded by the plug 166, leading to the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D. Supply cooling air. The plug 166 can be press-fit into the first primary cooling passage 152A or it can be fixedly coupled to the tip shroud 128.

  The first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D may have any suitable configuration and cross section. In the embodiment shown in FIG. 10, the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D are all linear. Nevertheless, in other embodiments, some or all of the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D may be curved. The first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D may have a circular cross-sectional shape or other suitable shape. The diameters of the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D are preferably uniform along their length, and the first primary cooling passage 152A It is smaller than the minimum diameter (eg, diameter 168 (FIG. 5)). Nevertheless, the diameters of the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D are non-uniform and are the same as the minimum diameter of the first primary cooling passage 152A. It can be the same or larger.

  The first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D can be formed in any suitable manner. For example, the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D can be formed by drilling or machining the rotor blade 100 after casting. In other embodiments, the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D are formed during the casting process using, for example, a jacketed core 182 (FIG. 9). can do.

  The first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D provide film cooling and / or convection cooling to the rotor blade 100. More specifically, the cooling air flows through the first primary cooling passage 152A, thereby absorbing the heat from the rotor blade 100 by convection as described in detail above. Plug 166 prevents cooling air from exiting first primary cooling passage 152A through first primary cooling passage outlet 154A. In this way, the first primary cooling passage 152A acts as a plenum when closed by the plug 166, leading to the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D. Supply cooling air. Cooling air then passes through the corresponding first, second, third, and fourth outlets 164A, 164B, 164C, 164D to the first, second, third, and fourth secondary cooling passages. Exit from 162A, 162B, 162C, 162D. While the cooling air flows through the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D, heat from the rotor blade 100 can be absorbed by convection. The first and second secondary cooling passages 162A, 162B can direct cooling air to the radially outer surface 142 of the tip shroud 128 and form a film of cooling air thereon. The third and fourth secondary cooling passages 162C, 162D can direct cooling air to the outer surface 180 of the transition portion 150 and form a film of cooling air thereon. The film of cooling air acts as a barrier to insulate the rotor blade 100 from the combustion gas 34 flowing through the hot gas path 32. In some embodiments, the secondary cooling passage may not be oriented to allow film cooling.

  As described in detail above, the first and second primary cooling passages 152A, 152B cool the airfoil 126, tip shroud 128, and transition portion 150 by convection. More specifically, the pressure side and / or suction side surfaces 146, 148 of the tip shroud 128 define first and second primary cooling passage outlets 154A, 154B. In this regard, the first and second primary cooling passages 152A, 152B increase the cooling of the tip shroud 128 compared to conventional cooling passage configurations. Further, the first and second primary cooling passages 152A, 152B are preferably proximate and disposed along the curvature of the outer surface 180 of the transition portion 150 of the rotor blade 100. As such, the first and second primary cooling passages 152A, 152B also improve cooling of the transition portion 150 over conventional cooling passage configurations. In certain embodiments, the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D provide membrane cooling and / or additional convective cooling to the tip shroud 128 and transition portion 150. Providing, thereby providing additional cooling to the rotor blade 100. Further, the first and second primary cooling passages 152A, 152B and / or the first, second, third, and fourth secondary cooling passages 162A, 162B, 162C, 162D are different from the conventional cooling passage configuration. In contrast, it does not undesirably increase the weight of the rotor blade 100.

This specification uses examples to disclose the technology, including the best mode. Also, examples are used to enable any person skilled in the art to practice the technology, including making and using any device or system and performing any integrated method. The patentable scope of the technology is defined by the claims, and may include other examples that occur to those skilled in the art. Where such other embodiments include structural elements that do not differ from the literal wording of the claims, or where they include equivalent structural elements that do not substantially differ from the literal wording of the claims. Such other embodiments are intended to be within the scope of the claims.
[Embodiment 1]
A rotor blade (28, 100) for a gas turbine engine (10) comprising:
An airfoil (126);
A tip shroud (128) having side surfaces (146, 148) and a radially outer surface (106, 142);
A transition portion (150) that couples the tip shroud (128) to the airfoil (126);
The airfoil (126), the transition portion (150), and the tip shroud (128) collectively define a primary cooling passageway (152) therein,
The rotor blade (28), wherein the primary cooling passage (152) includes a primary cooling passage outlet (154) defined by side surfaces (146, 148) of the tip shroud (128).
[Embodiment 2]
The rotor blade (28, 100) according to embodiment 1, wherein a pressure side (146) of the tip shroud (128) defines the primary cooling passage outlet (154).
[Embodiment 3]
2. The rotor blade (28, 100) according to embodiment 1, wherein a suction side surface (148) of the tip shroud (128) defines the primary cooling passage outlet (154).
[Embodiment 4]
2. The rotor blade (28, 100) of embodiment 1, wherein the tip shroud (128) defines a secondary cooling passage (162) in fluid communication with the primary cooling passage (152).
[Embodiment 5]
The secondary cooling passage (162) extends from the primary cooling passage (152) to a secondary cooling passage outlet (164) defined by a radially outer surface (106, 142) of the tip shroud (128). The rotor blade (28, 100) according to embodiment 4.
[Embodiment 6]
It further includes a plug (166) disposed at the primary cooling passage outlet (154), the plug (166) being adapted to direct the fluid flow through the secondary cooling passage (162) to the primary cooling passage outlet (154). The rotor blade (28, 100) according to embodiment 4, wherein the fluid flow through) is occluded.
[Embodiment 7]
Embodiment 6. The rotor blade (28, 100) of embodiment 4, wherein the secondary cooling passage (162) is straight.
[Embodiment 8]
The tip shroud (128) and the transition portion (150) collectively define a secondary cooling passage (162), and the secondary cooling passage (162) extends from the primary cooling passage (152) to the transition. The rotor blade (28, 100) of embodiment 1, extending to a secondary cooling passage outlet (164) defined by the outer surface of the portion (150).
[Embodiment 9]
2. The rotor blade (28, 100) of embodiment 1, wherein a portion of the primary cooling passageway (152) defined by the transition portion (150) of the rotor blade (28, 100) is curvilinear.
[Embodiment 10]
A portion of the primary cooling passage (152) defined by the airfoil (126) and a portion of the primary cooling passage (152) defined by the tip shroud (128) are linear. Embodiment 10. The rotor blade (28, 100) according to embodiment 9.
[Embodiment 11]
A portion of the primary cooling passage (152) defined by the airfoil (126) and a portion of the primary cooling passage (152) defined by the tip shroud (128) are curvilinear. Embodiment 10. The rotor blade (28, 100) according to embodiment 9.
[Embodiment 12]
2. The rotor blade (28, 100) of embodiment 1, wherein the primary cooling passage (152) includes one or more turbulators.
[Embodiment 13]
The rotor blade (28, 100) according to embodiment 1, wherein the primary cooling passage outlet (154) has an oval cross section.
[Embodiment 14]
2. The rotor blade of embodiment 1, wherein the airfoil (126), the transition portion (150), and the tip shroud (128) collectively define a plurality of primary cooling passages (152) therein. (28,100).
[Embodiment 15]
A gas turbine engine (10) comprising:
A compressor part (14);
A combustion part (16);
A turbine portion (18) including one or more rotor blades (28, 100), each rotor blade (28, 100) comprising:
An airfoil (126);
A tip shroud (128) including side surfaces (146, 148) and a radially outer surface (106, 142);
A transition portion (150) that couples the tip shroud (128) to the airfoil (126);
The airfoil (126), the transition portion (150), and the tip shroud (128) collectively define a primary cooling passageway (152) therein,
The gas turbine engine (10), wherein the primary cooling passage (152) includes a primary cooling passage outlet (154) defined by side surfaces (146, 148) of the tip shroud (128).
[Embodiment 16]
Further included is a secondary cooling passage (162) extending from the primary cooling passage (152) to a secondary cooling passage outlet (164) defined by a radially outer surface (106, 142) of the tip shroud (128). A gas turbine engine (10) according to embodiment 15,.
[Embodiment 17]
It further includes a plug (166) disposed at the primary cooling passage outlet (154), the plug (166) being adapted to direct the fluid flow through the secondary cooling passage (162) to the primary cooling passage outlet (154). The gas turbine engine (10) of embodiment 16, wherein the fluid flow through the
[Embodiment 18]
The tip shroud (128) and the transition portion (150) collectively define a secondary cooling passage (162), and the secondary cooling passage (162) extends from the primary cooling passage (152) to the transition. The gas turbine engine (10) of embodiment 15, extending to a secondary cooling passage outlet (164) defined by an outer surface of the portion (150).
[Embodiment 19]
The gas turbine engine (10) of embodiment 15, wherein a portion of the primary cooling passageway (152) defined by the transition portion (150) is curvilinear.
[Embodiment 20]
A portion of the primary cooling passage (152) defined by the airfoil (126) and a portion of the primary cooling passage (152) defined by the tip shroud (128) are linear. A gas turbine engine (10) according to embodiment 19.

DESCRIPTION OF SYMBOLS 10 Gas turbine engine 12 Inlet part 14 Compressor part 16 Combustion part 18 Turbine part 20 Exhaust part 22 Shaft 24 Rotor shaft 26 Rotor disk 28 Rotor blade 30 Outer casing 32 Hot gas path 34 Combustion gas 100 Rotor blade 102 Platform 104 Radius of platform Directional inner surface 106 platform radially outer surface 108 platform leading edge surface 110 platform trailing edge surface 112 platform pressure side slash surface 114 platform suction side slash surface 116 shank portion 118 angel wing 120 shank pocket 122 root portion 124 intake port 126 Airfoil 128 Tip shroud 130 Airfoil root 132 Airfoil span 134 Airfoil pressure side wall 136 Airfoil Pressure wall 138 Leading edge 140 Trailing edge 142 Radial outer surface 144 of tip shroud Rail 146 Pressure side surface 148 of tip shroud Pressure side surface 150 of tip shroud Transition portion 152 Primary cooling passage 152A First primary cooling passage 152B Second primary cooling passage Cooling passage 154A First primary cooling passage outlet 154B Second primary cooling passage outlet 156A Cooling passage airfoil portion 156B Cooling passage airfoil portion 158A First cooling passage transition portion 158B Second cooling passage transition portion 160A Cooling passage tip shroud portion 160B Cooling passage tip shroud portion 162A First secondary cooling passage 162B Second secondary cooling passage 162C Third secondary cooling passage 162D Fourth secondary cooling passage 164A First Secondary cooling passage outlet 164B Second secondary cooling passage outlet 164C Third two Cooling passage outlet 164D Fourth secondary cooling passage outlet 166 Plug 168 First diameter 170 Second diameter 172 Rectangular protrusion 174 Hemispherical protrusion 176 Dimple 178 Radial inner surface 180 of tip shroud Outer surface 182 of transition portion Core 184 with jacket Ring Sleeve 186 Core material 188 Mold 190 Melted material

Claims (15)

A rotor blade (28, 100) for a gas turbine engine (10) comprising:
An airfoil (126);
A tip shroud (128) having side surfaces (146, 148) and a radially outer surface (106, 142);
A transition portion (150) that couples the tip shroud (128) to the airfoil (126);
The airfoil (126), the transition portion (150), and the tip shroud (128) collectively define a primary cooling passageway (152) therein,
The rotor blade (28), wherein the primary cooling passage (152) includes a primary cooling passage outlet (154) defined by side surfaces (146, 148) of the tip shroud (128).   The rotor blade (28, 100) of claim 1, wherein a pressure side (146) of the tip shroud (128) defines the primary cooling passage outlet (154).   The rotor blade (28, 100) of claim 1, wherein a suction side (148) of the tip shroud (128) defines the primary cooling passage outlet (154).   The rotor blade (28, 100) of claim 1, wherein the tip shroud (128) defines a secondary cooling passage (162) in fluid communication with the primary cooling passage (152).   The secondary cooling passage (162) extends from the primary cooling passage (152) to a secondary cooling passage outlet (164) defined by a radially outer surface (106, 142) of the tip shroud (128). The rotor blade (28, 100) according to claim 4.   It further includes a plug (166) disposed at the primary cooling passage outlet (154), the plug (166) being adapted to direct the fluid flow through the secondary cooling passage (162) to the primary cooling passage outlet (154). The rotor blade (28, 100) according to claim 4, wherein the fluid flow through   The rotor blade (28, 100) according to claim 4, wherein the secondary cooling passage (162) is straight.   The tip shroud (128) and the transition portion (150) collectively define a secondary cooling passage (162), and the secondary cooling passage (162) extends from the primary cooling passage (152) to the transition. The rotor blade (28, 100) of claim 1, extending to a secondary cooling passage outlet (164) defined by an outer surface of the portion (150).   The rotor blade (28, 100) of claim 1, wherein a portion of the primary cooling passageway (152) defined by the transition portion (150) of the rotor blade (28, 100) is curved.   A portion of the primary cooling passage (152) defined by the airfoil (126) and a portion of the primary cooling passage (152) defined by the tip shroud (128) are linear. The rotor blade (28, 100) according to claim 9.   A portion of the primary cooling passage (152) defined by the airfoil (126) and a portion of the primary cooling passage (152) defined by the tip shroud (128) are curvilinear. The rotor blade (28, 100) according to claim 9.   The rotor blade (28, 100) of any preceding claim, wherein the primary cooling passage (152) includes one or more turbulators.   The rotor blade (28, 100) of claim 1, wherein the primary cooling passage outlet (154) has an oval cross section.   The rotor blade of claim 1, wherein the airfoil (126), the transition portion (150), and the tip shroud (128) collectively define a plurality of primary cooling passages (152) therein. (28,100). A gas turbine engine (10) comprising:
A compressor part (14);
A combustion part (16);
A turbine portion (18) including one or more rotor blades (28, 100), each rotor blade (28, 100) comprising:
An airfoil (126);
A tip shroud (128) including side surfaces (146, 148) and a radially outer surface (106, 142);
A transition portion (150) that couples the tip shroud (128) to the airfoil (126);
The airfoil (126), the transition portion (150), and the tip shroud (128) collectively define a primary cooling passageway (152) therein,
The gas turbine engine (10), wherein the primary cooling passage (152) includes a primary cooling passage outlet (154) defined by side surfaces (146, 148) of the tip shroud (128).